Gas turbine engine for long range aircraft

ABSTRACT

A gas turbine engine comprises a fan for delivering air into a bypass duct as bypass flow, into a core housing as core flow, with the core housing containing an upstream compressor rotor and a downstream compressor rotor. An overall pressure ratio is defined across the upstream and downstream compressor rotors. A bypass ratio is defined as a volume of air delivered as bypass flow compared to a volume of air delivered into the core housing. The overall pressure ratio is greater than or equal to about 45.0, and the bypass ratio is greater than or equal to about 11.0.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No. 14/107,273, filed Dec. 16, 2013.

BACKGROUND OF THE INVENTION

This application relates to a gas turbine engine designed for use on longer range aircraft.

Gas turbine engines are known and may include a fan delivering air into a bypass duct as propulsion air. In addition, the fan typically delivers air into a core housing and to a compressor. There are, typically, at least two compressor rotors with an upstream or lower pressure rotor compressing the air and then delivering it into a downstream or higher pressure rotor. The compressed air from the downstream compressor rotor is typically delivered into a combustion section where it is mixed with fuel and ignited. Products of this combustion may pass downstream over turbine rotors including an upstream turbine rotor that drives the downstream compressor rotor and a downstream turbine rotor that drives the upstream compressor rotor.

In one type engine, the downstream turbine rotor also drove a fan rotor, such that the fan rotor, the upstream compressor rotor, and the downstream turbine rotor all rotated at a single speed. More recently, a gear reduction has been placed between the fan rotor and the downstream turbine rotor or the fan drive turbine.

It is desirable to increase the compression ratio or the amount of compression done to air across the two compressor rotors. However, there has been a significant limitation in that the stress and temperature at the downstream end of the downstream compressor rotor limits how high the overall compression ratio may reach.

This area must be designed to withstand the repeated application of the highest stress situations for the gas turbine engine which typically occurs during take-off.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine comprises a fan for delivering air into a bypass duct as bypass flow, and into a core housing as core flow, with the core housing containing an upstream compressor rotor and a downstream compressor rotor. An overall pressure ratio is defined across the upstream and downstream compressor rotors. A bypass ratio is defined as a volume of air delivered as bypass flow compared to a volume of air delivered into the core housing. The overall pressure ratio is greater than or equal to about 45.0, and the bypass ratio is greater than or equal to about 11.0

In another embodiment according to the previous embodiment, the gas turbine engine is designed for use on long range aircraft defined as aircraft with at least two passenger aisles.

In another embodiment according to any of the previous embodiments, a fan drive turbine is configured to drive the upstream compressor rotor and the fan rotor through a gear reduction, with the fan drive turbine having at least three stages.

In another embodiment according to any of the previous embodiments, the fan drive turbine has six or fewer stages.

In another embodiment according to any of the previous embodiments, a ratio of a tip speed at the downstream compressor rotor compared to a tip speed at the upstream compressor rotor is less than or equal to about 1.18 and greater than or equal to about 1.0.

In another embodiment according to any of the previous embodiments, the bypass ratio is greater than or equal to about 14.0.

In another embodiment according to any of the previous embodiments, the overall pressure ratio across the upstream and downstream compressor rotors is equal to or greater than about 60.0.

In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal about 2.6.

In another embodiment according to any of the previous embodiments, the gas turbine engine is designed for use on long range aircraft defined as aircraft with a flight length equal to, or greater than about 6.0 hours.

In another embodiment according to any of the previous embodiments, the upstream compressor rotor has at least three stages.

In another embodiment according to any of the previous embodiments, the upstream compressor rotor has seven or fewer stages.

In another embodiment according to any of the previous embodiments, a fan drive turbine is configured to drive the upstream compressor rotor and the fan rotor through a gear reduction, with the fan drive turbine having at least three stages.

In another embodiment according to any of the previous embodiments, the fan drive turbine has six or fewer stages.

In another embodiment according to any of the previous embodiments, a ratio of a tip speed at the downstream compressor rotor compared to a tip speed at the upstream compressor rotor is less than or equal to about 1.18 and greater than or equal to about 1.0.

In another embodiment according to any of the previous embodiments, the bypass ratio is greater than or equal to about 14.0.

In another embodiment according to any of the previous embodiments, the overall pressure ratio across the upstream and downstream compressor rotors is equal to or greater than about 60.0.

In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal about 2.6.

In another embodiment according to any of the previous embodiments, a fan drive turbine is configured to drive the fan rotor through a gear reduction.

In another embodiment according to any of the previous embodiments, the fan drive turbine also is configured to drive the upstream compressor rotor, along with the fan rotor.

In another embodiment according to any of the previous embodiments, there are two additional turbine rotors upstream of the fan drive turbine for respectively driving the upstream and downstream compressor rotors.

In another embodiment according to any of the previous embodiments, a ratio of a tip speed at the downstream compressor rotor compared to a tip speed at the upstream compressor rotor is less than or equal to about 1.18 and greater than or equal to about 1.0.

In another embodiment according to any of the previous embodiments, the overall pressure ratio across the upstream and downstream compressor rotors is equal to or greater than about 60.0.

In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal about 2.6.

These and other features may be best understood from the following drawing and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A schematically shows a gas turbine engine wherein the fan-driving turbine also drives the upstream compressor.

FIG. 1B schematically shows an aircraft that may incorporate the engine of FIG. 1A or FIG. 2.

FIG. 2 schematically shows another gas turbine engine.

DETAILED DESCRIPTION

A gas turbine engine 20 shown schematically in FIG. 1A, is designed for use on long range aircraft.

As known, a fan rotor 22 delivers bypass air B within a nacelle 24. A core engine housing 26 receives core air flow C from the fan rotor 22. The core air flow C initially reaches an upstream compressor rotor 28, which compresses the air to a first lower level and then delivers that air into a downstream compressor rotor 38, where additional compression occurs.

The air from the compressor rotor 38 is delivered into a combustion section 42, mixed with fuel and ignited. Products of this combustion pass downstream over an upstream turbine rotor 40, which operates at a higher pressure and speed and drives a shaft 36 to drive the downstream compressor rotor 38. Downstream of the turbine rotor 40, the products of combustion drive a fan drive turbine 34 which is a downstream turbine rotor and operates at a lower pressure and speed than does the turbine rotor 40. The fan drive turbine 34 drives the upstream compressor rotor 28 through a shaft 30 and drives the fan rotor 22 through a gear reduction 32.

Applicant has recognized that on longer range aircraft the percentage of engine operation time at high stress level, such as take-off, becomes a very small percentage of the overall operation time. Long range aircraft may be defined as traveling 3,000 to 8,000 miles or more, and from about 6 to about 16 hours of flying time during a typical flight.

Often these aircraft can also be described as “twin aisle” aircraft because they are wide body aircraft as opposed to single aisle aircraft which are used on inter-continental flights or to feed airports that are used as hub-and-spoke airports for connecting flights.

As shown in FIG. 1B, an aircraft 100 schematically includes two engines 20. There are aisles 104 and 106 separating seating areas 102 and 108.

Take-off and climb will typically occur for only 45 seconds at take-off power followed by perhaps 20 minutes at climb. In such an example, the ratio of time at low power cruise to climb and take-off is at least 15 to 1 and can become as high as 40 to 1.

With such systems, it is possible to achieve higher pressure ratios because the high stress situations on the downstream end of the downstream compressor rotor 38 will occur much less frequently and over a small percentage of the overall operational time across the life of the engine.

In disclosed embodiments, the overall pressure ratio or amount of compression by the combined rotors 28 and 38 to the core air flow C may be equal to or above about 45.0, above about 60.0, and above about 65.0.

In addition, a bypass ratio or ratio of the volume of air delivered as bypass flow B to the volume of air delivered as core air flow C may be equal to or greater than about 11.0, and greater than or equal to 12.0, or greater than or equal to 14.0.

A gear ratio of the gear reduction 32 may be greater than or equal to about 2.6, greater than or equal to about 2.9, or greater than or equal to about 3.6.

The upstream compressor rotor 28 may have less than or equal to seven stages, or from three to seven stages. The fan drive turbine 34 may have three or more stages, or from three to six stages.

A tip speed at the tip 39 of the downstream compressor rotor 38 compared to the tip speed at tip 29 of the upstream compressor rotor 28 is less than or equal to about 1.8 and greater than or equal to about 1.0. Tip speed is defined as the tangential velocity of the leading edge of the longest blade.

Finally, the take-off thrust for the long range aircraft is greater than about 50,000 lbf at static CSLTO 86° F. The take-off thrust may be as high as 124,000 lbf at sea-level static condition.

Another factor for peak efficiency which this compression section is designed to produce is that there be at least two turbine rotors (40) in front of the fan-driving turbine. This allows for reasonable mach numbers through the turbine section ahead of the fan drive turbine which improves overall engine efficiency and reduces temperatures into the fan drive turbine to a manageable level. In contrast, a single turbine rotor ahead of the fan drive turbine may lower overall engine efficiency and drive up temperatures into the fan drive turbine to such an extent that at least one of the fan drive turbine stages may have to be cooled, owing to the stresses there. Cooling the fan drive turbine also lowers the efficiency of the overall engine cycle.

The fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1). The low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.

Another embodiment engine 120 is illustrated in FIG. 2. Many components are effectively identical to those shown in FIG. 1A, and carry the same numbers, simply increased by 100.

Whereas the FIG. 1A engine 20 has a fan drive turbine 34 driving a compressor 28, and a fan rotor 22 through the gear reduction 32, the FIG. 2 engine includes a higher pressure first turbine 140 driving a higher pressure downstream compressor section 138. An intermediate turbine section 141 drives the first compressor rotor, or the upstream compressor rotor 128. The fan drive turbine 134 only drives the fan rotor 122 through a gear reduction 132.

The quantities as described above with regard to the FIG. 1A embodiment, would also be true of the FIG. 2 embodiment.

Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention. 

What is claimed is:
 1. A gas turbine engine comprising: a fan for delivering air into a bypass duct as bypass flow and into a core housing as core flow; an upstream compressor rotor and a downstream compressor rotor; an overall pressure ratio greater than or equal to 45.0; a bypass ratio defined as a volume of air delivered as bypass flow compared to a volume of air delivered into said core housing, and said bypass ratio greater than or equal to 11; a fan drive turbine driving said upstream compressor rotor and driving said fan through a gear reduction, said fan drive turbine having at least three stages; and wherein a ratio of a tip speed at said downstream compressor rotor compared to a tip speed at said upstream compressor rotor is less than or equal to 1.18.
 2. The gas turbine engine as set forth in claim 1, further comprising two additional turbine rotors upstream of said fan drive turbine.
 3. The gas turbine engine as set forth in claim 2, wherein said fan drive turbine has six or fewer stages.
 4. The gas turbine engine as set forth in claim 3, wherein said ratio of said tip speed at said downstream compressor rotor compared to said tip speed at said upstream compressor rotor is greater than or equal to 1.0.
 5. The gas turbine engine as set forth in claim 4, wherein said fan drive turbine having fewer stages than said upstream compressor rotor.
 6. The gas turbine engine as set forth in claim 4, wherein said fan has a low fan pressure ratio of less than 1.45.
 7. The gas turbine engine as set forth in claim 6, wherein said gas turbine engine being designed for use on a long range aircraft, said long range aircraft being defined at least by one of an aircraft with at least two passenger aisles, or an aircraft with a flight length equal to, or greater than 6.0 hours.
 8. The gas turbine engine as set forth in claim 7, wherein said fan drive turbine has a turbine pressure ratio that is greater than five 5:1, said turbine pressure ratio defined as a pressure measured prior to an inlet of said fan drive turbine as related to a pressure at an outlet of said fan drive turbine prior to an exhaust nozzle.
 9. The gas turbine engine as set forth in claim 6, wherein said fan drive turbine has a turbine pressure ratio that is greater than five 5:1, said turbine pressure ratio defined as a pressure measured prior to an inlet of said fan drive turbine as related to a pressure at an outlet of said fan drive turbine prior to an exhaust nozzle.
 10. The gas turbine engine as set forth in claim 9, wherein: said upstream compressor rotor has seven or fewer stages; a gear ratio of said gear reduction is greater than or equal 2.6; and said bypass ratio is greater than or equal to 14.0.
 11. The gas turbine engine as set forth in claim 10, wherein said gear reduction is a planetary gear system and said overall pressure ratio is equal to or greater than
 60. 12. The gas turbine engine as set forth in claim 11, wherein said upstream compressor rotor has at least three stages.
 13. The gas turbine engine as set forth in claim 12, further comprising two turbine rotors upstream of said fan drive turbine.
 14. The gas turbine engine as set forth in claim 3, wherein said fan has a low fan pressure ratio of less than 1.45.
 15. The gas turbine engine as set forth in claim 14, wherein said fan drive turbine has a turbine pressure ratio that is greater than five 5:1, said turbine pressure ratio defined as a pressure measured prior to an inlet of said fan drive turbine as related to a pressure at an outlet of said fan drive turbine prior to an exhaust nozzle.
 16. The gas turbine engine as set forth in claim 15, wherein: said fan drive turbine has six or fewer stages; and said upstream compressor rotor has seven or fewer stages.
 17. The gas turbine engine as set forth in claim 16, wherein a gear ratio of said gear reduction is greater than or equal 2.6, and said bypass ratio is greater than or equal to 14.0.
 18. The gas turbine engine as set forth in claim 17, wherein said ratio of a tip speed at said downstream compressor rotor compared to said tip speed at said upstream compressor rotor is greater than or equal to 1.0.
 19. The gas turbine engine as set forth in claim 18, wherein said gas turbine engine being designed for use on a long range aircraft, said long range aircraft being defined at least by one of an aircraft with at least two passenger aisles, or an aircraft with a flight length equal to, or greater than 6.0 hours.
 20. The gas turbine engine as set forth in claim 18, wherein: said gear reduction is a planetary gear system; a gear ratio of said gear reduction is greater than or equal 3.6; and said overall pressure ratio is equal to or greater than
 65. 21. The gas turbine engine as set forth in claim 20, wherein said gas turbine engine being designed for use on a long range aircraft, said long range aircraft being defined at least by one of an aircraft with at least two passenger aisles, or an aircraft with a flight length equal to, or greater than 6.0 hours.
 22. A method of designing a gas turbine engine for a long range aircraft comprising: driving a fan through a gear reduction with a fan drive turbine, with said fan delivering air into a bypass duct as bypass flow and into a core housing as core flow, with said core flow flowing into an upstream compressor rotor and a downstream compressor rotor, wherein a bypass ratio is defined as a volume of air delivered as bypass flow compared to a volume of air delivered into said core housing, said bypass ratio greater than or equal to 11, said fan drive turbine having at least three stages but no more than six stages, and two additional turbine rotors upstream of said fan drive turbine; driving said upstream and downstream compressor rotors, with an overall pressure ratio greater than or equal to 45.0; and wherein said gas turbine engine being mounted on said long range aircraft during operation, with said long range aircraft being defined at least by one of an aircraft with at least two passenger aisles, or an aircraft with a flight length equal to, or greater than 6.0 hours.
 23. The method as set forth in claim 22, wherein: the step of driving said upstream and downstream compressor rotors establishes a ratio of a tip speed at said downstream compressor rotor compared to a tip speed at said upstream compressor rotor that is less than or equal to 1.18 and is greater than or equal to 1.0; and the step of driving said fan occurs for a duration of said flight length, said flight length including a first time at a low power cruise condition and a second time at a climb condition and a take-off condition, and a ratio of said first time to said second time being at least 15:1.
 24. The method as set forth in claim 23, wherein: said fan drive turbine has fewer than six stages; and said upstream compressor rotor has at least three stages, but fewer than seven stages.
 25. The method as set forth in claim 24, wherein said long range aircraft is defined as an aircraft with a flight length of 3,000 to 8,000 miles.
 26. The method as set forth in claim 25, wherein said gas turbine engine is a first engine and a second engine, each of said first and second engines mounted on said long range aircraft during operation such that a take-off thrust of said long range aircraft is greater than 50,000 lbf at static CSLTO 86° F.
 27. The method as set forth in claim 22, wherein: the step of driving said upstream and downstream compressor rotors includes driving said upstream compressor rotor with said fan drive turbine; said bypass ratio is greater than or equal to 14.0; said fan has a low fan pressure ratio of less than 1.45; and said fan drive turbine has a turbine pressure ratio that is greater than five 5:1, said turbine pressure ratio defined as a pressure measured prior to an inlet of said fan drive turbine as related to a pressure at an outlet of said fan drive turbine prior to an exhaust nozzle.
 28. The method as set forth in claim 27, wherein said overall pressure ratio equal to or greater than
 60. 29. The method as set forth in claim 28, wherein said gas turbine engine is a first engine and a second engine, each of said first and second engines mounted on said long range aircraft during operation such that a take-off thrust of said long range aircraft is greater than 50,000 lbf at static CSLTO 86° F., and said step of driving said upstream and downstream compressor rotors establishes a ratio of a tip speed at said downstream compressor rotor compared to a tip speed at said upstream compressor rotor that is less than or equal to 1.18.
 30. The method as set forth in claim 27, wherein the step of driving said upstream and downstream compressor rotors establishes a ratio of a tip speed at said downstream compressor rotor compared to a tip speed at said upstream compressor rotor that is greater than or equal to 1.0. 